Many satellites and other spacecraft are powered from one or more solar arrays. The solar arrays are typically held in a retracted position during launch, and are then moved to a deployed position following, or during portions of, the spacecraft orbit initialization. In addition to the solar arrays, many spacecraft include one or more energy storage flywheels to provide both a backup power source and to provide attitude control for the vehicle. In such systems, each flywheel is controlled and regulated to balance the electrical demand in the vehicle electrical distribution system, and is also controlled in response to programmed or remote attitude (or torque) commands received by the vehicle main controller.
Some satellites and other spacecraft may additionally be implemented with one or more relatively high power loads. Such spacecraft, which are generally referred to herein as high power spacecraft, are generally safe and reliable. However, high power spacecraft can present challenging issues with respect to electrical power distribution bus voltage and orbit initialization. Each of these issues will be briefly discussed, beginning with the issue associated with electrical power distribution bus voltage.
Standard spacecraft components are typically designed to operate at a relatively low voltage, and are thus supplied with electrical power from a relatively low voltage power distribution bus. For example, many standard spacecraft are implemented with a 28 VDC power distribution bus. However, it is relatively inefficient to use a relatively low voltage power distribution bus to supply high power loads, due to the large associated current. One proposed solution to this drawback is to increase the voltage of the power distribution bus. The increased voltage allows smaller currents to provide the same amount of power to the high power loads. The smaller currents in turn result in lower line losses, and increased power distribution system efficiency. However, this solution presents its own drawback, in that most of the components on a spacecraft are typically low voltage loads (e.g., 28 VDC), and will thus need a voltage regulator, or other similar device. As is generally known, such devices also exhibit characteristic inefficiencies, which can eliminate any advantage that high voltage power distribution provides.
As regards orbit initialization, spacecraft have historically included onboard chemical batteries, which are fully charged when the spacecraft and its associated launch vehicle are launched. Typically, a spacecraft is launched in a “turned-off” state, and is then switched to a “turned-on” state when it is ejected from its associated launch vehicle. The batteries are sized to provide sufficient power to conduct the spacecraft orbit initialization process, prior to the spacecraft's solar arrays being deployed to collect energy from the sun. However, the amount of power needed to implement orbit initialization can result in an undesirably large battery size, which can increase overall spacecraft weight and cost.
Hence, there is a need for a system and method of providing power generation and attitude control for a high power spacecraft that addresses one or more of the above-noted drawbacks. Namely, a system and method that can efficiently supply electrical power to both low voltage and high voltage loads and/or a system and method that does not rely on relatively heavy batteries to supply power during spacecraft orbit initialization. The present invention addresses at least these needs.